US4254618A - Cooling air cooler for a gas turbofan engine - Google Patents
Cooling air cooler for a gas turbofan engine Download PDFInfo
- Publication number
- US4254618A US4254618A US05/825,614 US82561477A US4254618A US 4254618 A US4254618 A US 4254618A US 82561477 A US82561477 A US 82561477A US 4254618 A US4254618 A US 4254618A
- Authority
- US
- United States
- Prior art keywords
- air
- turbine
- cooling air
- cooling
- engine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Lifetime
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
- F02C7/185—Cooling means for reducing the temperature of the cooling air or gas
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F28—HEAT EXCHANGE IN GENERAL
- F28D—HEAT-EXCHANGE APPARATUS, NOT PROVIDED FOR IN ANOTHER SUBCLASS, IN WHICH THE HEAT-EXCHANGE MEDIA DO NOT COME INTO DIRECT CONTACT
- F28D21/00—Heat-exchange apparatus not covered by any of the groups F28D1/00 - F28D20/00
- F28D21/0001—Recuperative heat exchangers
- F28D21/0014—Recuperative heat exchangers the heat being recuperated from waste air or from vapors
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
- B64D2033/024—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes comprising cooling means
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F28—HEAT EXCHANGE IN GENERAL
- F28D—HEAT-EXCHANGE APPARATUS, NOT PROVIDED FOR IN ANOTHER SUBCLASS, IN WHICH THE HEAT-EXCHANGE MEDIA DO NOT COME INTO DIRECT CONTACT
- F28D21/00—Heat-exchange apparatus not covered by any of the groups F28D1/00 - F28D20/00
- F28D2021/0019—Other heat exchangers for particular applications; Heat exchange systems not otherwise provided for
- F28D2021/0021—Other heat exchangers for particular applications; Heat exchange systems not otherwise provided for for aircrafts or cosmonautics
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- This invention relates to gas turbines and, more particularly, to a concept for efficiently reducing the temperature of air used to cool high temperature turbines in gas turbofan engines.
- Modern aircraft gas turbofan engines operate at turbine inlet air temperature levels which are beyond the structural temperature capabilities of high temperature alloys. Hence, engine hot flow path components and, in particular, turbine blades and vanes must be cooled in order to assure their structural integrity in order to meet operating life requirements. It is well understood that gas turbine engine shaft horsepower and specific fuel consumption (which is the rate of fuel consumption per unit of power output) can be improved by increasing turbine inlet temperature. In order to take advantage of this potential performance improvement, modern turbine cooling technology utilizes air-cooled, hollow turbine nozzle vanes and blades to permit operation at inlet gas temperatures in excess of 2000° F. (1094° C.). In general, these sophisticated methods of turbine cooling have utilized compressor discharge or interstage bleed air as a coolant.
- the benefits obtained from sophisticated air-cooling techniques are at least partially offset by the extraction of the necessary cooling air from the propulsive cycle.
- the cooling airflow rate required is a function of the hot gas temperature, increasing with increasing hot gas temperature.
- the compressor bleed air used for cooling must bypass the combustor and one or more turbine stages, thus giving rise to a performance penalty proportionate to the amount of cooling air utilized.
- the air that is bled from the compressor and used as cooling air for the turbine rotor blades has had work done on it by the compressor.
- it because it is normally returned into the flow path gas stream downstream of the turbine nozzle, it does not return its full measure of work to the cycle as it expands through the turbine.
- engine performance can be increased by reducing the amount of cooling air required by the turbine. Reducing the cooling airflow rate results in improved engine performance with a consequent reduction in specific fuel consumption, the actual magnitude of the cooling airflow rate and specific fuel consumption reductions which can be realized being a function of the specific engine application.
- One method of reducing the amount of cooling air required by the turbine is to cool the cooling air entering the hot components.
- One widely advocated method of cooling the cooling air is to utilize the heat sink capability available in the engine fuel.
- the relatively hot cooling air is placed in heat exchange relationship with the relatively cool engine fuel, thereby cooling the cooling air and heating the fuel.
- the energy extracted by the fuel is reintroduced back into the propulsive cycle as the heated fuel is burned in the combustor, thereby producing what has commonly been referred to as a "regenerative engine".
- the above objects are obtained in an aircraft gas turbofan engine by providing a heat exchanger wherein the turbine cooling air and relatively cooler air from the fan bypass duct are maintained in heat exchange relationship, thereby cooling the turbine cooling air.
- the turbine cooling air is bled, for example, from the discharge of the compressor through ports in the engine casing at various circumferential locations and is ducted to the heat exchanger which is disposed inwardly of the fan bypass portion of the gas turbine engine.
- the relatively cool fan bypass duct air is bled at the inner wall of the fan duct into a diffuser where the dynamic head of the fan stream is largely recovered.
- the fan bleed air is then ducted through the heat exchanger into heat exchange relationship with the relatively warmer compressor discharge bleed air, thereby absorbing heat from the cooling air, and returned to the fan bypass duct.
- the cooled compressor discharge bleed air is then routed to the high pressure turbine through the compressor rear frame struts and is expanded through an expander nozzle prior to cooling the high pressure turbine components.
- the cooling flow rates through the heat exchanger may be reduced by increasing the magnitude of the cooling air temperature reduction in the heat exchanger in direct proportion to the reduction in flow rates.
- the resulting over-cooled cooling air is then mixed with uncooled compressor discharge bleed air ahead of the expander nozzle to obtain the cooling air temperature reduction necessary to cool the turbine.
- FIG. 1 is a simplified cross-sectional view, in partial cutaway, of an aircraft gas turbofan incorporating the preferred embodiment of the subject invention and illustrating the relationship of the heat exchanger to the various other engine components;
- FIG. 2 is a simplified cross-sectional view of a portion of the gas turbofan engine of FIG. 1 depicting an alternative embodiment of the cooling system of the present invention
- FIG. 3 graphically depicts the turbine relative cooling flow rate and specific fuel consumption reductions as a function of the change in cooling air temperature for the representative gas turbofan engine of FIG. 1.
- FIG. 1 a representative gas turbofan engine designated generally at 10, and which incorporates the present invention, is diagrammatically shown. While it is recognized that turbofan engines are, by now, well known in the art, a brief description of the operation of the engine will enhance appreciation of the interrelationship of the various components in light of the invention soon to be described.
- this engine may be considered as comprising a core engine 12, a fan 14 including a rotatable stage of fan blades 16 (only one of which is shown for clarity), and a fan turbine (not shown) downstream of the core engine in the area generally depicted as 17 and which is interconnected to the fan 14 by shaft 18.
- the core engine 12 includes an axial flow compressor 20 having a rotor 22. Air enters inlet 24 from the left of FIG. 1 and is initially compressed by the fan blades 16. A first portion of this relatively cool compressed air enters the fan bypass duct 26 defined, in part, by core engine 12 and a circumscribing fan cowl or nacelle 28 and discharges through a fan nozzle 30. A second portion of the compressed air enters core engine inlet 32, is further compressed by the axial flow compressor 20 and is discharged to a combustor 34 where it is mixed with fuel and burned to provide high energy combustion gases which drive a core engine turbine 36. The turbine 36, in turn, drives the rotor 22 through a shaft 38 in the usual manner of a gas turbine engine.
- gas turbine engine shaft horsepower and specific fuel consumption (which is the rate of fuel consumption per unit of power output) can be improved by increasing the temperature at the inlet to the core engine turbine 36 (sometimes referred to as the "high pressure turbine”).
- turbine 36 since modern aircraft turbofan engines operate at turbine inlet air temperature levels which are beyond the structural temperature capabilities of high temperature alloys, turbine 36 must be cooled to assure its structural integrity. It can, therefore, be appreciated that as the temperature of the hot exhaust gases exiting combustor 34 is increased, an increased percentage of cooling air is required to cool the turbine.
- the source of the coolant for the turbine 36 has been air bled from the discharge of compressor 20 which is routed to and through the turbine in a manner well known in the art.
- the compressor discharge has been the logical choice for the coolant flow since the pressure of the compressor discharge airflow (referred to hereinafter as the "cooling air") is high enough to drive the cooling air through the tortuous path associated with the turbine structure.
- the cooling air has had work performed on it by the compressor, its temperature level has been increased.
- compressor compression ratios are increased, and as aircraft velocities increase, a corresponding rise in the temperature of the cooling air is experienced.
- an increasingly higher percentage of cooling flow is required to cool the turbine to acceptable temperature levels.
- this cooling air must bypass the combustor and perhaps one or more turbine stages before being returned to the propulsive cycle, thus giving rise to a performance penalty in proportion to the amount of cooling air used. It thus becomes advantageous to reduce the amount of cooling air required.
- FIG. 3 there is depicted in graphical form the change in turbine relative cooling flow rates and specific fuel consumption as a function of the change in cooling air temperature for a typical gas turbofan engine of the variety depicted in FIG. 1.
- FIG. 3 an estimate of the cooling airflow and specific fuel consumption reductions that can be realized by cooling the turbine blade cooling air of a two-stage core engine turbine of current design is shown in FIG. 3. It may be observed from the figure that in this particular application, reducing the cooling air temperature by 250° F. results in a 50 percent reduction in the required cooling airflow rate with a corresponding reduction of 1.1 percent in specific fuel consumption. It is clear from this simplistic example that great benefits can be obtained by reducing the temperature of the turbine cooling air.
- the present invention contemplates the use of the relatively cool fan bypass stream as a heat sink to cool the cooling air.
- the engine is provided with a means for capturing a portion of the relatively cool bypass flow such as, for example, shroud 44 which circumscribes a portion of the length of core engine 12 within the bypass duct to define a flow passage 46 (perhaps in the form of an annulus) therebetween.
- a heat exchanger 54 Disposed within this passage is a heat exchanger 54, preferably of the cross-flow tubular type which is described in greater particularity in the copending patent applications of Thomas G. Wakeman, Ser. No. 849,139, filed on Oct. 14, 1977, issued on Oct. 17, 1978, as U.S. Pat. No. 4,020,150, and Ser.
- Turbine cooling air is bled from the compressor discharge through ports 48 in the core engine casing 50 at various circumferential locations and routed through at least one conduit 52 to the heat exchanger 54.
- the bypass air portion captured by shroud 44 enters a diffuser section 55 where the dynamic head of the captured portion is largely recovered and ducted through the heat exchanger where it absorbs heat from the turbine cooling air.
- This bypass air portion is then returned to the fan duct at the discharge 56 of passage 46.
- the cooling air thus cooled is routed via conduit 58 to the high pressure turbine 36 through compressor rear frame struts 60 and thereafter to the expander nozzle 62 of a type taught by U.S. Pat. No. 3,565,545, issued to Melvin Bobo et al, which is assigned to the assignee of the present invention.
- the cooled cooling air then travels via passageway 64 to turbine 36 where it is used to perform the cooling function in a manner well known in the art.
- FIG. 1 may be modified as in FIG. 2 by reducing the cooling flow rate ducted through heat exchanger 54 and increasing the magnitude of the cooling air temperature reduction in direct proportion to the reduction in flow rate. While this design approach reduces the size of the required ducting 58, it will generally result in some increase in heat exchanger weight in order to increase the effectiveness of the heat exchanger. In such an embodiment, an auxiliary hot flow of cooling air is bled from the core engine through apertures 66 and 68 in core engine inner casing structure 70.
- This uncooled auxiliary bleed air is mixed with the cooled cooling air exiting the downstream end 72 of conduit 58 ahead of the expander nozzle 62 to obtain the desired final cooling air temperature.
- the resulting mixture is then utilized to cool the hot turbine components as in FIG. 1 and in accordance with well known turbine cooling principles.
- the present invention is readily adaptable to existing gas turbofan engines in that the components may be designed and placed in the engine in such a manner that they do not substantially change the configuration or design of nearby existing structure.
- the heat exchanger is of the air-to-air variety and is completely independent of the need for highly volatile coolant fluid which characterize prior art turbine cooling concepts.
- the present invention contemplates the cooling of the turbine coolant by placing it in heat exchange relationship with the abundant supply of fan bypass air in the gas turbofan engine
- the particular configuration of the heat exchanger may take many forms, such as heat exchangers of the single or multiple-pass variety.
- it may be desirable to extract the cooling air from the compressor 20 at a location other than the compressor discharge.
- the present invention may be used to cool the cooling air required for any of a number of high temperature turbine components and is not limited to cooling the cooling air required for turbine blades and vanes. It is intended that the appended claims cover all such variations in the present invention's broader inventive concepts.
Abstract
An air-to-air heat exchanger is provided for a gas turbofan engine to significantly reduce the quantity of cooling air that is presently needed to effectively cool the hot turbine parts. Typically, the turbine is internally cooled with air bled from the compressor which, though cooler than the turbine, has been heated due to the work done on it by the compressor. In accordance with the present invention, the heat exchanger is located internally of the bypass duct to place in heat exchange relationship a captured portion of the relatively cool bypass flow and this warmer compressor bleed air, thereby cooling the turbine coolant and significantly reducing the amount of such coolant required. This results in a decrease in engine specific fuel consumption.
Description
This invention relates to gas turbines and, more particularly, to a concept for efficiently reducing the temperature of air used to cool high temperature turbines in gas turbofan engines.
Modern aircraft gas turbofan engines operate at turbine inlet air temperature levels which are beyond the structural temperature capabilities of high temperature alloys. Hence, engine hot flow path components and, in particular, turbine blades and vanes must be cooled in order to assure their structural integrity in order to meet operating life requirements. It is well understood that gas turbine engine shaft horsepower and specific fuel consumption (which is the rate of fuel consumption per unit of power output) can be improved by increasing turbine inlet temperature. In order to take advantage of this potential performance improvement, modern turbine cooling technology utilizes air-cooled, hollow turbine nozzle vanes and blades to permit operation at inlet gas temperatures in excess of 2000° F. (1094° C.). In general, these sophisticated methods of turbine cooling have utilized compressor discharge or interstage bleed air as a coolant. However, the benefits obtained from sophisticated air-cooling techniques are at least partially offset by the extraction of the necessary cooling air from the propulsive cycle. It can be appreciated that the cooling airflow rate required is a function of the hot gas temperature, increasing with increasing hot gas temperature. Furthermore, the compressor bleed air used for cooling must bypass the combustor and one or more turbine stages, thus giving rise to a performance penalty proportionate to the amount of cooling air utilized. More particularly, the air that is bled from the compressor and used as cooling air for the turbine rotor blades has had work done on it by the compressor. However, because it is normally returned into the flow path gas stream downstream of the turbine nozzle, it does not return its full measure of work to the cycle as it expands through the turbine. Additionally, the reintroduction of cooling air into the hot gas stream produces a loss in gas stream total pressure. This is a result of the momentum mixing losses associated with injecting relatively low total pressure cooling air into a high total pressure gas stream. Thus, the greater the amount of cooling air which is routed through the turbine blades, the greater the losses associated with the coolant become on the propulsive cycle. Thus, while turbine blade cooling has inherent advantages, it also has associated therewith certain inherent disadvantages which are functions of the quantity of cooling air used in cooling the turbine rotor blades.
It will, therefore, be appreciated that engine performance can be increased by reducing the amount of cooling air required by the turbine. Reducing the cooling airflow rate results in improved engine performance with a consequent reduction in specific fuel consumption, the actual magnitude of the cooling airflow rate and specific fuel consumption reductions which can be realized being a function of the specific engine application.
One method of reducing the amount of cooling air required by the turbine is to cool the cooling air entering the hot components. One widely advocated method of cooling the cooling air is to utilize the heat sink capability available in the engine fuel. In such a scheme, the relatively hot cooling air is placed in heat exchange relationship with the relatively cool engine fuel, thereby cooling the cooling air and heating the fuel. The energy extracted by the fuel is reintroduced back into the propulsive cycle as the heated fuel is burned in the combustor, thereby producing what has commonly been referred to as a "regenerative engine". While various studies indicate that fuel-air heat exchangers offer an advantage of small size and low weight, the fuels currently used in aircraft engines (JP4, JP5) are limited in their heat sink capacity, the available heat sink already being used largely to cool the engine oil. To obtain an additional heat sink capacity to permit cooling of the cooling air would require the use of special fuels such as JP7 or JP9, which are currently unavailable in commercial quantities. Additionally, the use of fuel in a fuel-air heat exchanger presents a potential fire hazard which may be unacceptable for commercial engine applications. It will, therefore, be appreciated that another technique for cooling the cooling air is required in order to reduce the coolant flow rate and thereby enhance overall engine performance.
Accordingly, it is the primary object of the present invention to provide for a reduction in the amount of cooling air required by the turbine of a gas turbofan engine by reducing the temperature of the cooling air passing therethrough in order to enhance overall engine performance.
This, and other objects and advantages, will be more clearly understood from the following detailed descriptions, the drawing and specific examples, all of which are intended to be typical of, rather than in any way limiting on, the scope of the present invention.
Briefly stated, the above objects are obtained in an aircraft gas turbofan engine by providing a heat exchanger wherein the turbine cooling air and relatively cooler air from the fan bypass duct are maintained in heat exchange relationship, thereby cooling the turbine cooling air. The turbine cooling air is bled, for example, from the discharge of the compressor through ports in the engine casing at various circumferential locations and is ducted to the heat exchanger which is disposed inwardly of the fan bypass portion of the gas turbine engine. The relatively cool fan bypass duct air is bled at the inner wall of the fan duct into a diffuser where the dynamic head of the fan stream is largely recovered. The fan bleed air is then ducted through the heat exchanger into heat exchange relationship with the relatively warmer compressor discharge bleed air, thereby absorbing heat from the cooling air, and returned to the fan bypass duct. The cooled compressor discharge bleed air is then routed to the high pressure turbine through the compressor rear frame struts and is expanded through an expander nozzle prior to cooling the high pressure turbine components. In an alternative embodiment of the present invention where the space available for ducting the cooling air through the compressor rear frame struts is limited, the cooling flow rates through the heat exchanger may be reduced by increasing the magnitude of the cooling air temperature reduction in the heat exchanger in direct proportion to the reduction in flow rates. The resulting over-cooled cooling air is then mixed with uncooled compressor discharge bleed air ahead of the expander nozzle to obtain the cooling air temperature reduction necessary to cool the turbine.
Incorporation of this heat exchanger into an aircraft gas turbofan engine permits a reduction in the quantity of air required for turbine cooling and, thus, provides an improvement in engine performance. Conversely, an increase in blade life can be achieved by maintaining the original coolant flow rate but by reducing the temperature of the coolant, with essentially no further degradation in engine performance.
While the specification concludes with claims particularly pointing out and distinctly claiming the subject matter which is regarded as part of the present invention, it is believed that the invention will be more fully understood from the following description of the preferred embodiments which are given by way of example with the accompanying drawing in which:
FIG. 1 is a simplified cross-sectional view, in partial cutaway, of an aircraft gas turbofan incorporating the preferred embodiment of the subject invention and illustrating the relationship of the heat exchanger to the various other engine components;
FIG. 2 is a simplified cross-sectional view of a portion of the gas turbofan engine of FIG. 1 depicting an alternative embodiment of the cooling system of the present invention; and
FIG. 3 graphically depicts the turbine relative cooling flow rate and specific fuel consumption reductions as a function of the change in cooling air temperature for the representative gas turbofan engine of FIG. 1.
Referring to the drawing wherein like numerals correspond to like elements throughout, attention is first directed to FIG. 1 wherein a representative gas turbofan engine designated generally at 10, and which incorporates the present invention, is diagrammatically shown. While it is recognized that turbofan engines are, by now, well known in the art, a brief description of the operation of the engine will enhance appreciation of the interrelationship of the various components in light of the invention soon to be described. Basically, this engine may be considered as comprising a core engine 12, a fan 14 including a rotatable stage of fan blades 16 (only one of which is shown for clarity), and a fan turbine (not shown) downstream of the core engine in the area generally depicted as 17 and which is interconnected to the fan 14 by shaft 18. The core engine 12 includes an axial flow compressor 20 having a rotor 22. Air enters inlet 24 from the left of FIG. 1 and is initially compressed by the fan blades 16. A first portion of this relatively cool compressed air enters the fan bypass duct 26 defined, in part, by core engine 12 and a circumscribing fan cowl or nacelle 28 and discharges through a fan nozzle 30. A second portion of the compressed air enters core engine inlet 32, is further compressed by the axial flow compressor 20 and is discharged to a combustor 34 where it is mixed with fuel and burned to provide high energy combustion gases which drive a core engine turbine 36. The turbine 36, in turn, drives the rotor 22 through a shaft 38 in the usual manner of a gas turbine engine. The hot gases of combustion then pass through and drive the fan turbine which, in turn, drives the fan 14. A propulsive force is thus obtained by the action of the fan 14 discharging air from the fan bypass duct 26 through the fan nozzle 30 and by the discharge of combustion gases from a core engine nozzle 40 defined, in part, by plug 42. The above description is typical of many present-day gas turbofan engines and is not meant to be limiting to the present invention, as it will become readily apparent from the following description that the present invention is capable of application to any gas turbofan engine of the bypass variety and is not necessarily restricted to use with the particular configuration depicted herein. The foregoing description of the operation of the engine depicted in FIG. 1 is, therefore, merely meant to be illustrative of one type of application.
It is also well understood that gas turbine engine shaft horsepower and specific fuel consumption (which is the rate of fuel consumption per unit of power output) can be improved by increasing the temperature at the inlet to the core engine turbine 36 (sometimes referred to as the "high pressure turbine"). However, since modern aircraft turbofan engines operate at turbine inlet air temperature levels which are beyond the structural temperature capabilities of high temperature alloys, turbine 36 must be cooled to assure its structural integrity. It can, therefore, be appreciated that as the temperature of the hot exhaust gases exiting combustor 34 is increased, an increased percentage of cooling air is required to cool the turbine. Traditionally, the source of the coolant for the turbine 36 has been air bled from the discharge of compressor 20 which is routed to and through the turbine in a manner well known in the art. The compressor discharge has been the logical choice for the coolant flow since the pressure of the compressor discharge airflow (referred to hereinafter as the "cooling air") is high enough to drive the cooling air through the tortuous path associated with the turbine structure. However, because the cooling air has had work performed on it by the compressor, its temperature level has been increased. And, as compressor compression ratios are increased, and as aircraft velocities increase, a corresponding rise in the temperature of the cooling air is experienced. As a result, an increasingly higher percentage of cooling flow is required to cool the turbine to acceptable temperature levels. As mentioned earlier, this cooling air must bypass the combustor and perhaps one or more turbine stages before being returned to the propulsive cycle, thus giving rise to a performance penalty in proportion to the amount of cooling air used. It thus becomes advantageous to reduce the amount of cooling air required.
Referring now to FIG. 3 there is depicted in graphical form the change in turbine relative cooling flow rates and specific fuel consumption as a function of the change in cooling air temperature for a typical gas turbofan engine of the variety depicted in FIG. 1. As an illustration, an estimate of the cooling airflow and specific fuel consumption reductions that can be realized by cooling the turbine blade cooling air of a two-stage core engine turbine of current design is shown in FIG. 3. It may be observed from the figure that in this particular application, reducing the cooling air temperature by 250° F. results in a 50 percent reduction in the required cooling airflow rate with a corresponding reduction of 1.1 percent in specific fuel consumption. It is clear from this simplistic example that great benefits can be obtained by reducing the temperature of the turbine cooling air.
The present invention contemplates the use of the relatively cool fan bypass stream as a heat sink to cool the cooling air. Referring again to FIG. 1, it may be seen that the engine is provided with a means for capturing a portion of the relatively cool bypass flow such as, for example, shroud 44 which circumscribes a portion of the length of core engine 12 within the bypass duct to define a flow passage 46 (perhaps in the form of an annulus) therebetween. Disposed within this passage is a heat exchanger 54, preferably of the cross-flow tubular type which is described in greater particularity in the copending patent applications of Thomas G. Wakeman, Ser. No. 849,139, filed on Oct. 14, 1977, issued on Oct. 17, 1978, as U.S. Pat. No. 4,020,150, and Ser. No. 797,669 filed on May 17, 1977. Turbine cooling air is bled from the compressor discharge through ports 48 in the core engine casing 50 at various circumferential locations and routed through at least one conduit 52 to the heat exchanger 54. The bypass air portion captured by shroud 44 enters a diffuser section 55 where the dynamic head of the captured portion is largely recovered and ducted through the heat exchanger where it absorbs heat from the turbine cooling air. This bypass air portion is then returned to the fan duct at the discharge 56 of passage 46. The cooling air thus cooled is routed via conduit 58 to the high pressure turbine 36 through compressor rear frame struts 60 and thereafter to the expander nozzle 62 of a type taught by U.S. Pat. No. 3,565,545, issued to Melvin Bobo et al, which is assigned to the assignee of the present invention. The cooled cooling air then travels via passageway 64 to turbine 36 where it is used to perform the cooling function in a manner well known in the art.
In order to permit the efficient return of the heated bypass flow portion back into bypass duct 26 upstream of fan nozzle 30, its static pressure must be matched to the static pressure in the bypass duct at location 56 where the bleed portion is reintroduced. Thus, the total pressure drop of the bled portion, including the pressure drop through the diffuser section 55, heat exchanger 54 and flow passage 46 must be limited to a value less than or equal to the dynamic head of the remainder of the bypass flow stream at the location where the bled portion is reintroduced into the fan duct.
If, as is the case in existing gas turbofan engines for which the present invention may wish to be adapted, the space available for ducting through the compressor rear frame struts 60 is limited, the configuration of FIG. 1 may be modified as in FIG. 2 by reducing the cooling flow rate ducted through heat exchanger 54 and increasing the magnitude of the cooling air temperature reduction in direct proportion to the reduction in flow rate. While this design approach reduces the size of the required ducting 58, it will generally result in some increase in heat exchanger weight in order to increase the effectiveness of the heat exchanger. In such an embodiment, an auxiliary hot flow of cooling air is bled from the core engine through apertures 66 and 68 in core engine inner casing structure 70. This uncooled auxiliary bleed air is mixed with the cooled cooling air exiting the downstream end 72 of conduit 58 ahead of the expander nozzle 62 to obtain the desired final cooling air temperature. The resulting mixture is then utilized to cool the hot turbine components as in FIG. 1 and in accordance with well known turbine cooling principles.
It thus becomes clear from the foregoing descriptions that the stated objects of the present invention have been attained in the embodiments as depicted and that an engine configured in accordance with the present invention will have significant performance benefits over prior art gas turbofan engines. In particular, reliance has been placed on the well-established concept of utilizing compressor bleed air as a turbine blade coolant. However, the amount of compressor bleed air required has been substantially reduced, thereby enhancing overall cycle performance. Conversely, an increase in blade life can be achieved by maintaining the original coolant flow rate but by reducing the temperature of the coolant, with essentially no further degradation in engine performance. Furthermore, the present invention is readily adaptable to existing gas turbofan engines in that the components may be designed and placed in the engine in such a manner that they do not substantially change the configuration or design of nearby existing structure. Furthermore, the heat exchanger is of the air-to-air variety and is completely independent of the need for highly volatile coolant fluid which characterize prior art turbine cooling concepts.
It should become obvious to one skilled in the art that certain changes can be made to the above-described invention without departing from the broad inventive concepts thereof. For example, while the present invention contemplates the cooling of the turbine coolant by placing it in heat exchange relationship with the abundant supply of fan bypass air in the gas turbofan engine, the particular configuration of the heat exchanger may take many forms, such as heat exchangers of the single or multiple-pass variety. Furthermore, it may be desirable to extract the cooling air from the compressor 20 at a location other than the compressor discharge. In addition, the present invention may be used to cool the cooling air required for any of a number of high temperature turbine components and is not limited to cooling the cooling air required for turbine blades and vanes. It is intended that the appended claims cover all such variations in the present invention's broader inventive concepts.
Claims (2)
1. In a gas turbine engine of the bypass variety having a fan for pressurizing a cool flow of fan bypass air; a core engine including a compressor for pressurizing a hot flow of cooling air; a turbine of the air cooled variety; a heat exchanger for receiving a portion of the cool bypass air and a portion of the hot cooling air, wherein heat is transferred directly from the hot cooling air portion to the bypass air portion thereby resulting in a flow of cooled cooling air; and a means for routing the cooled cooling air to the turbine for turbine cooling; an improvement comprising:
a duct for routing the bypass air around the core engine;
means for capturing a portion of the bypass air from said duct, said capturing means defining a flow passage having a diffuser section to substantially recover the dynamic head of the fan bypass portion, and wherein said heat exchanger is disposed within said passage downstream of said diffuser section.
2. The improved gas turbine engine, as recited in claim 1, wherein said capturing means defining said flow passage returns the captured portion of the fan bypass air to said duct upstream of a fan nozzle.
Priority Applications (5)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US05/825,614 US4254618A (en) | 1977-08-18 | 1977-08-18 | Cooling air cooler for a gas turbofan engine |
FR7823116A FR2400618A1 (en) | 1977-08-18 | 1978-08-04 | AIR-COOLED BLOWER TURBO ENGINE AND PROCESS FOR COOLING THIS ENGINE |
IT26716/78A IT1098088B (en) | 1977-08-18 | 1978-08-11 | COOLING AIR COOLER FOR TURBO-FAN GAS ENGINE |
DE19782835903 DE2835903A1 (en) | 1977-08-18 | 1978-08-16 | COOLING AIR COOLER FOR A GAS TURBINE ENGINE |
JP10009078A JPS5452216A (en) | 1977-08-18 | 1978-08-18 | Method of and apparatus for cooling gas turbine engine |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US05/825,614 US4254618A (en) | 1977-08-18 | 1977-08-18 | Cooling air cooler for a gas turbofan engine |
Publications (1)
Publication Number | Publication Date |
---|---|
US4254618A true US4254618A (en) | 1981-03-10 |
Family
ID=25244478
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US05/825,614 Expired - Lifetime US4254618A (en) | 1977-08-18 | 1977-08-18 | Cooling air cooler for a gas turbofan engine |
Country Status (5)
Country | Link |
---|---|
US (1) | US4254618A (en) |
JP (1) | JPS5452216A (en) |
DE (1) | DE2835903A1 (en) |
FR (1) | FR2400618A1 (en) |
IT (1) | IT1098088B (en) |
Cited By (170)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4542623A (en) * | 1983-12-23 | 1985-09-24 | United Technologies Corporation | Air cooler for providing buffer air to a bearing compartment |
US4546605A (en) * | 1983-12-16 | 1985-10-15 | United Technologies Corporation | Heat exchange system |
US4561246A (en) * | 1983-12-23 | 1985-12-31 | United Technologies Corporation | Bearing compartment for a gas turbine engine |
US4574584A (en) * | 1983-12-23 | 1986-03-11 | United Technologies Corporation | Method of operation for a gas turbine engine |
US4601202A (en) * | 1983-12-27 | 1986-07-22 | General Electric Company | Gas turbine engine component cooling system |
US4608819A (en) * | 1983-12-27 | 1986-09-02 | General Electric Company | Gas turbine engine component cooling system |
US4709545A (en) * | 1983-05-31 | 1987-12-01 | United Technologies Corporation | Bearing compartment protection system |
US4791782A (en) * | 1986-08-27 | 1988-12-20 | Rolls-Royce Plc | Fluid outlet duct |
WO1990001624A1 (en) * | 1988-08-09 | 1990-02-22 | Sundstrand Corporation | High pressure intercooled turbine engine |
US4991394A (en) * | 1989-04-03 | 1991-02-12 | Allied-Signal Inc. | High performance turbine engine |
US5012646A (en) * | 1988-11-28 | 1991-05-07 | Machen, Inc. | Turbine engine having combustor air precooler |
US5038560A (en) * | 1988-10-22 | 1991-08-13 | Rolls-Royce Plc | Fluid outlet duct |
US5163285A (en) * | 1989-12-28 | 1992-11-17 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." | Cooling system for a gas turbine |
US5203163A (en) * | 1990-08-01 | 1993-04-20 | General Electric Company | Heat exchange arrangement in a gas turbine engine fan duct for cooling hot bleed air |
EP0564135A2 (en) * | 1992-03-23 | 1993-10-06 | General Electric Company | Gas turbine engine cooling system |
US5269133A (en) * | 1991-06-18 | 1993-12-14 | General Electric Company | Heat exchanger for cooling a gas turbine |
US5269135A (en) * | 1991-10-28 | 1993-12-14 | General Electric Company | Gas turbine engine fan cooled heat exchanger |
US5297386A (en) * | 1992-08-26 | 1994-03-29 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) | Cooling system for a gas turbine engine compressor |
US5319927A (en) * | 1992-07-18 | 1994-06-14 | Rolls-Royce Plc | Ducted fan gas turbine engine |
US5392614A (en) * | 1992-03-23 | 1995-02-28 | General Electric Company | Gas turbine engine cooling system |
US5394687A (en) * | 1993-12-03 | 1995-03-07 | The United States Of America As Represented By The Department Of Energy | Gas turbine vane cooling system |
US5581996A (en) * | 1995-08-16 | 1996-12-10 | General Electric Company | Method and apparatus for turbine cooling |
US5782077A (en) * | 1995-05-15 | 1998-07-21 | Aerospatiale Societe Nationale Industrielle | Device for bleeding off and cooling hot air in an aircraft engine |
US6058696A (en) * | 1997-12-22 | 2000-05-09 | United Technologies Corporation | Inlet and outlet module for a heat exchanger for a flowpath for working medium gases |
US6106229A (en) * | 1997-12-22 | 2000-08-22 | United Technologies Corporation | Heat exchanger system for a gas turbine engine |
US6134880A (en) * | 1997-12-31 | 2000-10-24 | Concepts Eti, Inc. | Turbine engine with intercooler in bypass air passage |
WO2001065095A1 (en) * | 2000-02-29 | 2001-09-07 | Mtu Aero Engines Gmbh | Cooling air system |
US20030147741A1 (en) * | 2000-05-15 | 2003-08-07 | Alessandro Coppola | Device for controlling the cooling flows of gas turbines |
EP1503061A1 (en) * | 2003-07-28 | 2005-02-02 | Snecma Moteurs | Method for cooling hot turbine parts using air partly cooled by an external heat exchanger and correspondingly cooled gas turbine engine |
US20050150970A1 (en) * | 2004-01-13 | 2005-07-14 | Snecma Moteurs | Cooling system for hot parts of an aircraft engine, and aircraft engine equipped with such a cooling system |
US20050268612A1 (en) * | 2004-04-24 | 2005-12-08 | Rolls-Royce Plc | Engine |
US20060042225A1 (en) * | 2004-08-27 | 2006-03-02 | Pratt & Whitney Canada Corp. | Bypass duct fluid cooler |
US20070245738A1 (en) * | 2006-04-20 | 2007-10-25 | Stretton Richard G | Heat exchanger arrangement |
US20070245739A1 (en) * | 2006-04-20 | 2007-10-25 | Stretton Richard G | Gas turbine engine |
US20070264133A1 (en) * | 2006-05-11 | 2007-11-15 | United Technologies Corporation | Thermal management system for turbofan engines |
US20080016845A1 (en) * | 2006-07-19 | 2008-01-24 | United Technologies Corporation | Lubricant cooling exchanger dual intake duct |
US20080053060A1 (en) * | 2006-08-29 | 2008-03-06 | Pratt & Whitney Canada Corp. | Bypass lip seal |
WO2008025136A1 (en) * | 2006-08-29 | 2008-03-06 | Pratt & Whitney Canada Corp. | Turbofan bypass duct air cooled fluid cooler installation |
US20080095611A1 (en) * | 2006-10-19 | 2008-04-24 | Michael Ralph Storage | Method and apparatus for operating gas turbine engine heat exchangers |
US20080141656A1 (en) * | 2006-08-01 | 2008-06-19 | Snecma | Bypass turbomachine with artificial variation of its throat section |
EP1967716A1 (en) | 2007-02-27 | 2008-09-10 | Snecma | Aircraft engine equipped with heat exchange system |
US20080271433A1 (en) * | 2007-05-03 | 2008-11-06 | Pratt & Whitney Canada Corp. | Low profile bleed air cooler |
US20080310955A1 (en) * | 2007-06-13 | 2008-12-18 | United Technologies Corporation | Hybrid cooling of a gas turbine engine |
US20080314047A1 (en) * | 2007-06-25 | 2008-12-25 | Honeywell International, Inc. | Cooling systems for use on aircraft |
EP2011988A2 (en) * | 2007-07-06 | 2009-01-07 | General Electric Company | Heat exchanger for a turbine engine |
US20090007567A1 (en) * | 2006-01-19 | 2009-01-08 | Airbus France | Dual Flow Turbine Engine Equipped with a Precooler |
US20090097972A1 (en) * | 2007-10-10 | 2009-04-16 | United Technologies Corp. | Gas Turbine Engine Systems and Related Methods Involving Heat Exchange |
ES2322317A1 (en) * | 2007-06-20 | 2009-06-18 | Futur Investment Partners, S.A. | Turbopropulsor aeronautical (Machine-translation by Google Translate, not legally binding) |
US20090169359A1 (en) * | 2007-12-26 | 2009-07-02 | Michael Joseph Murphy | Heat exchanger arrangement for turbine engine |
US20090188232A1 (en) * | 2008-01-28 | 2009-07-30 | Suciu Gabriel L | Thermal management system integrated pylon |
US20090301057A1 (en) * | 2006-06-27 | 2009-12-10 | Airbus France | Turboreactor for aircraft |
US20100043396A1 (en) * | 2008-08-25 | 2010-02-25 | General Electric Company | Gas turbine engine fan bleed heat exchanger system |
WO2010037450A2 (en) | 2008-10-03 | 2010-04-08 | Rolls-Royce Plc | Turbine cooling system |
US20100139288A1 (en) * | 2008-12-10 | 2010-06-10 | Pratt & Whitney Canada Corp. | Heat exchanger to cool turbine air cooling flow |
GB2468346A (en) * | 2009-03-06 | 2010-09-08 | Rolls Royce Plc | Cooling system for gas turbine engine |
US20100300066A1 (en) * | 2007-06-25 | 2010-12-02 | Airbus Operations (Societe Par Actions Simplifiee) | Turbojet engine for aircraft |
EP2275656A2 (en) | 2009-07-15 | 2011-01-19 | Rolls-Royce plc | System for cooling cooling-air in a gas turbine engine |
US20110011058A1 (en) * | 2009-07-17 | 2011-01-20 | Rolls-Royce Deutschland Ltd & Co Kg | Turbofan engine |
US20110088405A1 (en) * | 2009-10-15 | 2011-04-21 | John Biagio Turco | Gas turbine engine temperature modulated cooling flow |
US20110150634A1 (en) * | 2009-12-21 | 2011-06-23 | Denis Bajusz | Integration of an Air-Liquid Heat Exchanger on an Engine |
CN101374722B (en) * | 2006-01-19 | 2011-07-06 | 法国空中巴士公司 | Dual flow turbine engine equipped with a precooler |
EP1650407A3 (en) * | 2004-10-19 | 2011-07-27 | General Electric Company | Method and apparatus for cooling gas turbine engines |
WO2011139317A2 (en) | 2009-12-31 | 2011-11-10 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and heat exchange system |
EP2447543A1 (en) * | 2010-10-27 | 2012-05-02 | Siemens Aktiengesellschaft | Axial compressor and method for operating same |
US20120114467A1 (en) * | 2010-11-04 | 2012-05-10 | Elder James S | Gas turbine engine heat exchanger with tapered fins |
US20120144843A1 (en) * | 2009-12-31 | 2012-06-14 | Eric Sean Donovan | Gas turbine engine and cooling system |
US20120159961A1 (en) * | 2010-12-24 | 2012-06-28 | Michael Stephen Krautheim | Gas turbine engine heat exchanger |
US8266888B2 (en) * | 2010-06-24 | 2012-09-18 | Pratt & Whitney Canada Corp. | Cooler in nacelle with radial coolant |
US20120282079A1 (en) * | 2010-04-09 | 2012-11-08 | Glahn Jorn A | Rear hub cooling for high pressure compressor |
EP2527603A2 (en) | 2011-05-27 | 2012-11-28 | General Electric Company | Flade Duct Turbine Cooling And Power And Thermal Management |
US8397487B2 (en) | 2011-02-28 | 2013-03-19 | General Electric Company | Environmental control system supply precooler bypass |
US20130152602A1 (en) * | 2011-12-14 | 2013-06-20 | Rolls-Royce Plc | Controller |
US20130186102A1 (en) * | 2012-01-25 | 2013-07-25 | Honeywell International Inc. | Gas turbine engine in-board cooled cooling air system |
US20130239583A1 (en) * | 2012-03-14 | 2013-09-19 | United Technologies Corporation | Pump system for hpc eps parasitic loss elimination |
WO2013147953A1 (en) * | 2011-12-30 | 2013-10-03 | Rolls-Royce North American Technologies Inc. | Aircraft propulsion gas turbine engine with heat exchange |
US20130291554A1 (en) * | 2012-05-02 | 2013-11-07 | Pratt & Whitney Canada Corp. | Air cooler system for gas turbine engines |
US8602717B2 (en) | 2010-10-28 | 2013-12-10 | United Technologies Corporation | Compression system for turbomachine heat exchanger |
US20140010639A1 (en) * | 2012-07-05 | 2014-01-09 | Nathan Snape | Gas turbine engine oil tank with integrated packaging configuration |
US8636836B2 (en) | 2009-02-04 | 2014-01-28 | Purdue Research Foundation | Finned heat exchangers for metal hydride storage systems |
US20140096534A1 (en) * | 2012-10-04 | 2014-04-10 | United Technologies Corporation | Low Profile Compressor Bleed Air-Oil Coolers |
US20140131027A1 (en) * | 2012-11-09 | 2014-05-15 | Rolls-Royce Plc | Heat exchange arrangement |
US8778063B2 (en) | 2009-02-04 | 2014-07-15 | Purdue Research Foundation | Coiled and microchannel heat exchangers for metal hydride storage systems |
WO2015047533A1 (en) | 2013-09-24 | 2015-04-02 | United Technologies Corporation | Bypass duct heat exchanger placement |
RU2550224C1 (en) * | 2013-11-25 | 2015-05-10 | Российская Федерация, от имени которой выступает Министерство промышленности и торговли Российской Федерации (Минпромторг России) | Gas turbine engine |
US20150128561A1 (en) * | 2013-03-13 | 2015-05-14 | Rolls-Royce North American Technologies, Inc. | Three stream, variable area, vectorable nozzle |
US9045998B2 (en) | 2011-12-12 | 2015-06-02 | Honeywell International Inc. | System for directing air flow to a plurality of plena |
US9051943B2 (en) | 2010-11-04 | 2015-06-09 | Hamilton Sundstrand Corporation | Gas turbine engine heat exchanger fins with periodic gaps |
US20150192033A1 (en) * | 2012-07-19 | 2015-07-09 | Snecma | Cooling of an oil circuit of a turbomachine |
WO2015105552A1 (en) | 2014-01-07 | 2015-07-16 | United Technologies Corporation | Cross-stream heat exchanger |
US9109514B2 (en) | 2012-01-10 | 2015-08-18 | Hamilton Sundstrand Corporation | Air recovery system for precooler heat-exchanger |
US20150247462A1 (en) * | 2012-09-28 | 2015-09-03 | United Technologies Corporation | Gas turbine engine thermal management system for heat exchanger using bypass flow |
US20150300254A1 (en) * | 2014-04-17 | 2015-10-22 | Rolls-Royce Plc | Propulsion engine |
EP2952698A1 (en) * | 2014-06-06 | 2015-12-09 | United Technologies Corporation | Turbine stage cooling |
US9222411B2 (en) | 2011-12-21 | 2015-12-29 | General Electric Company | Bleed air and hot section component cooling air system and method |
US20160024964A1 (en) * | 2013-03-15 | 2016-01-28 | United Technologies Corporation | Gas turbine engine with air-oil cooler oil tank |
US20160032763A1 (en) * | 2013-03-14 | 2016-02-04 | United Technologies Corporation | Heatshield discourager seal for a gas turbine engine |
US9255492B2 (en) | 2011-12-14 | 2016-02-09 | Rolls-Royce Plc | Gas turbine engine having a multi-variable closed loop controller for regulating tip clearance |
US9267390B2 (en) | 2012-03-22 | 2016-02-23 | Honeywell International Inc. | Bi-metallic actuator for selectively controlling air flow between plena in a gas turbine engine |
US9267434B2 (en) | 2012-01-29 | 2016-02-23 | United Technologies Corporation | Heat exchanger |
US20160230658A1 (en) * | 2015-02-10 | 2016-08-11 | United Technologies Corporation | Gas turbine engine and an airflow control system |
US20160237908A1 (en) * | 2015-02-12 | 2016-08-18 | United Technologies Corporation | Intercooled cooling air using existing heat exchanger |
US20160237905A1 (en) * | 2015-02-12 | 2016-08-18 | United Technologies Corporation | Intercooled cooling air |
US20160237907A1 (en) * | 2015-02-12 | 2016-08-18 | United Technologies Corporation | Intercooled cooling air with auxiliary compressor control |
US9422063B2 (en) | 2013-05-31 | 2016-08-23 | General Electric Company | Cooled cooling air system for a gas turbine |
US9429072B2 (en) | 2013-05-22 | 2016-08-30 | General Electric Company | Return fluid air cooler system for turbine cooling with optional power extraction |
US9562475B2 (en) | 2012-12-19 | 2017-02-07 | Siemens Aktiengesellschaft | Vane carrier temperature control system in a gas turbine engine |
US20170082028A1 (en) * | 2015-02-12 | 2017-03-23 | United Technologies Corporation | Intercooled cooling air using existing heat exchanger |
US20170167378A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam Via Turbine Extraction and Compressor Extraction |
US20170167376A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam Via Turbine Extraction |
US20170167377A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam Via Turbine Extraction |
US20170167379A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam via Turbine Extraction and Compressor Extraction |
US20170167375A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | Power Plant With Steam Generation Via Combustor Gas Extraction |
US20170184027A1 (en) * | 2015-12-29 | 2017-06-29 | General Electric Company | Method and system for compressor and turbine cooling |
RU2623854C1 (en) * | 2016-07-06 | 2017-06-29 | Публичное акционерное общество "Научно-производственное объединение "Сатурн" | Method of greasing and cooling front support of the rotor of the gas turbine engine |
US20170307311A1 (en) * | 2016-04-26 | 2017-10-26 | United Technologies Corporation | Simple Heat Exchanger Using Super Alloy Materials for Challenging Applications |
US20180051630A1 (en) * | 2016-08-22 | 2018-02-22 | United Technologies Corporation | Heat Exchanger for Gas Turbine Engine with Support Damper Mounting |
US10066550B2 (en) | 2014-05-15 | 2018-09-04 | Rolls-Royce North American Technologies, Inc. | Fan by-pass duct for intercooled turbo fan engines |
US10100739B2 (en) | 2015-05-18 | 2018-10-16 | United Technologies Corporation | Cooled cooling air system for a gas turbine engine |
US10125684B2 (en) | 2015-12-29 | 2018-11-13 | Pratt & Whitney Canada Corp. | Surface cooler for aero engine |
US20190003392A1 (en) * | 2017-01-27 | 2019-01-03 | United Technologies Corporation | Thermal shield for gas turbine engine diffuser case |
RU2679573C1 (en) * | 2018-02-16 | 2019-02-11 | Валерий Николаевич Сиротин | Cooling system of bearings of gas turbine engine turbines |
US10221862B2 (en) | 2015-04-24 | 2019-03-05 | United Technologies Corporation | Intercooled cooling air tapped from plural locations |
RU187493U1 (en) * | 2018-05-28 | 2019-03-11 | Публичное Акционерное Общество "Одк-Сатурн" | HEAT EXCHANGER COOLING DEVICE |
CN109723558A (en) * | 2017-10-30 | 2019-05-07 | 通用电气公司 | Gas-turbine unit and its operating method including heat management system |
US20190145317A1 (en) * | 2017-11-14 | 2019-05-16 | Rolls-Royce Plc | Gas turbine engine having an air-oil heat exchanger |
US10371055B2 (en) | 2015-02-12 | 2019-08-06 | United Technologies Corporation | Intercooled cooling air using cooling compressor as starter |
US10385777B2 (en) * | 2012-10-01 | 2019-08-20 | United Technologies Corporation | Bifurcated inlet scoop for gas turbine engine |
US20190301301A1 (en) * | 2018-04-02 | 2019-10-03 | General Electric Company | Cooling structure for a turbomachinery component |
US10443508B2 (en) | 2015-12-14 | 2019-10-15 | United Technologies Corporation | Intercooled cooling air with auxiliary compressor control |
US20190323431A1 (en) * | 2018-04-19 | 2019-10-24 | United Technologies Corporation | Intercooled cooling air with advanced cooling system |
US10480419B2 (en) | 2015-04-24 | 2019-11-19 | United Technologies Corporation | Intercooled cooling air with plural heat exchangers |
US10550768B2 (en) | 2016-11-08 | 2020-02-04 | United Technologies Corporation | Intercooled cooled cooling integrated air cycle machine |
US10577964B2 (en) | 2017-03-31 | 2020-03-03 | United Technologies Corporation | Cooled cooling air for blade air seal through outer chamber |
US10669940B2 (en) | 2016-09-19 | 2020-06-02 | Raytheon Technologies Corporation | Gas turbine engine with intercooled cooling air and turbine drive |
US10677166B2 (en) | 2015-08-12 | 2020-06-09 | Rolls-Royce North American Technologies Inc. | Heat exchanger for a gas turbine engine propulsion system |
US10711640B2 (en) | 2017-04-11 | 2020-07-14 | Raytheon Technologies Corporation | Cooled cooling air to blade outer air seal passing through a static vane |
US10718233B2 (en) | 2018-06-19 | 2020-07-21 | Raytheon Technologies Corporation | Intercooled cooling air with low temperature bearing compartment air |
US10731560B2 (en) * | 2015-02-12 | 2020-08-04 | Raytheon Technologies Corporation | Intercooled cooling air |
US10738703B2 (en) | 2018-03-22 | 2020-08-11 | Raytheon Technologies Corporation | Intercooled cooling air with combined features |
CN111577466A (en) * | 2020-06-22 | 2020-08-25 | 中国航空发动机研究院 | Ice-proof bleed air preheating and turbine cooling bleed air precooling system for aircraft engine |
US10794288B2 (en) | 2015-07-07 | 2020-10-06 | Raytheon Technologies Corporation | Cooled cooling air system for a turbofan engine |
US10794290B2 (en) | 2016-11-08 | 2020-10-06 | Raytheon Technologies Corporation | Intercooled cooled cooling integrated air cycle machine |
US10830145B2 (en) | 2018-04-19 | 2020-11-10 | Raytheon Technologies Corporation | Intercooled cooling air fleet management system |
US10830148B2 (en) | 2015-04-24 | 2020-11-10 | Raytheon Technologies Corporation | Intercooled cooling air with dual pass heat exchanger |
WO2020234524A1 (en) | 2019-05-20 | 2020-11-26 | Safran | Optimized heat exchange system for a turbomachine |
WO2020234525A2 (en) | 2019-05-20 | 2020-11-26 | Safran | Optimised heat exchange system |
CN112228226A (en) * | 2020-10-16 | 2021-01-15 | 中国航发四川燃气涡轮研究院 | Aircraft engine turbine rotor cooling thermal management system |
US10961911B2 (en) | 2017-01-17 | 2021-03-30 | Raytheon Technologies Corporation | Injection cooled cooling air system for a gas turbine engine |
US10995673B2 (en) | 2017-01-19 | 2021-05-04 | Raytheon Technologies Corporation | Gas turbine engine with intercooled cooling air and dual towershaft accessory gearbox |
US11060484B2 (en) * | 2018-06-29 | 2021-07-13 | The Boeing Company | Nozzle wall for an air-breathing engine of a vehicle and method therefor |
US11066997B2 (en) * | 2016-12-14 | 2021-07-20 | Safran Aircraft Engines | Fluid circuit in a turbine engine |
US11078837B2 (en) * | 2019-02-06 | 2021-08-03 | Raytheon Technologies Corporation | Engine bleed air ducting into heat exchanger |
US11162417B2 (en) * | 2018-05-22 | 2021-11-02 | General Electric Company | Scoop inlet |
US11174816B2 (en) | 2019-02-25 | 2021-11-16 | Rolls-Royce Corporation | Bypass duct conformal heat exchanger array |
US11255268B2 (en) | 2018-07-31 | 2022-02-22 | Raytheon Technologies Corporation | Intercooled cooling air with selective pressure dump |
US11268444B2 (en) * | 2017-05-18 | 2022-03-08 | Raytheon Technologies Corporation | Turbine cooling arrangement |
US11300002B2 (en) | 2018-12-07 | 2022-04-12 | Pratt & Whitney Canada Corp. | Static take-off port |
US20220235705A1 (en) * | 2021-01-26 | 2022-07-28 | General Electric Company | Heat transfer system |
CN114876644A (en) * | 2022-05-09 | 2022-08-09 | 北京航空航天大学 | Periodic porous bearing support plate |
US11492971B2 (en) * | 2019-09-06 | 2022-11-08 | Raytheon Technologies Corporation | Turbine engine system with heat exchanger in bypassable secondary duct |
US11542026B2 (en) * | 2017-12-21 | 2023-01-03 | Safran Nacelles | Aircraft propulsion unit and method for ventilating an engine enclosure |
US11549393B2 (en) * | 2019-02-18 | 2023-01-10 | Safran Aero Boosters Sa | Air-oil heat exchanger |
US11591965B2 (en) | 2021-03-29 | 2023-02-28 | General Electric Company | Thermal management system for transferring heat between fluids |
US11674396B2 (en) | 2021-07-30 | 2023-06-13 | General Electric Company | Cooling air delivery assembly |
US11684974B2 (en) | 2014-10-21 | 2023-06-27 | Raytheon Technologies Corporation | Additive manufactured ducted heat exchanger system |
US11702958B2 (en) | 2021-09-23 | 2023-07-18 | General Electric Company | System and method of regulating thermal transport bus pressure |
US11725584B2 (en) | 2018-01-17 | 2023-08-15 | General Electric Company | Heat engine with heat exchanger |
US11788469B2 (en) | 2020-11-16 | 2023-10-17 | General Electric Company | Thermal management system for a gas turbine engine |
US11788470B2 (en) | 2021-03-01 | 2023-10-17 | General Electric Company | Gas turbine engine thermal management |
US11808210B2 (en) | 2015-02-12 | 2023-11-07 | Rtx Corporation | Intercooled cooling air with heat exchanger packaging |
Families Citing this family (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE3514352A1 (en) * | 1985-04-20 | 1986-10-23 | MTU Motoren- und Turbinen-Union München GmbH, 8000 München | GAS TURBINE ENGINE WITH DEVICES FOR DIVERSING COMPRESSOR AIR FOR COOLING HOT PARTS |
DE3916477A1 (en) * | 1989-05-20 | 1990-11-22 | Mak Maschinenbau Krupp | Removing fuel from injection nozzle - involves releasing compressed air from storage vessel |
US5144794A (en) * | 1989-08-25 | 1992-09-08 | Hitachi, Ltd. | Gas turbine engine with cooling of turbine blades |
CA2046797A1 (en) * | 1990-08-01 | 1992-02-02 | Franklin D. Parsons | Heat exchange arrangement in a gas turbine engine fan duct for cooling hot bleed air |
DE10122695A1 (en) * | 2001-05-10 | 2002-11-21 | Siemens Ag | Process for cooling a gas turbine and gas turbine plant |
DE602006019008D1 (en) | 2006-10-12 | 2011-01-27 | United Technologies Corp | MODULATION FLOW THROUGH TURBO MOTOR COOLING SYSTEM |
US7823389B2 (en) * | 2006-11-15 | 2010-11-02 | General Electric Company | Compound clearance control engine |
FR2920483B1 (en) * | 2007-08-30 | 2009-10-30 | Snecma Sa | GENERATION OF ELECTRICITY IN A TURBOMACHINE |
DE102009011924A1 (en) * | 2009-03-10 | 2010-09-16 | Rolls-Royce Deutschland Ltd & Co Kg | Bypass duct of a turbofan engine |
DE102011106961A1 (en) * | 2011-07-08 | 2013-01-10 | Rolls-Royce Deutschland Ltd & Co Kg | Flight gas turbine engine i.e. turbomachine, has flow guide element designed as radiator element, and core thruster surrounded by by-pass channel, where partial flow is conducted from channel through engine casing for cooling core thruster |
DE102011106965A1 (en) * | 2011-07-08 | 2013-01-10 | Rolls-Royce Deutschland Ltd & Co Kg | Aircraft gas turbine engine with heat exchanger in the core engine housing |
US20140027097A1 (en) * | 2012-07-30 | 2014-01-30 | Ian Alexandre Araujo De Barros | Heat Exchanger for an Intercooler and Water Extraction Apparatus |
JP5844503B1 (en) * | 2014-10-21 | 2016-01-20 | 住友精密工業株式会社 | Heat exchanger for aircraft engine |
Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3034298A (en) * | 1958-06-12 | 1962-05-15 | Gen Motors Corp | Turbine cooling system |
US3224194A (en) * | 1963-06-26 | 1965-12-21 | Curtiss Wright Corp | Gas turbine engine |
US3528250A (en) * | 1969-04-16 | 1970-09-15 | Gen Motors Corp | Bypass engine with afterburning and compressor bleed air heat exchanger in bypass duct |
US3584458A (en) * | 1969-11-25 | 1971-06-15 | Gen Motors Corp | Turbine cooling |
US3651645A (en) * | 1969-10-11 | 1972-03-28 | Mtu Muenchen Gmbh | Gas turbine for aircrafts |
US3797561A (en) * | 1970-10-02 | 1974-03-19 | Secr Defence | Oil tanks and coolers |
US3842597A (en) * | 1973-03-16 | 1974-10-22 | Gen Electric | Gas turbine engine with means for reducing the formation and emission of nitrogen oxides |
-
1977
- 1977-08-18 US US05/825,614 patent/US4254618A/en not_active Expired - Lifetime
-
1978
- 1978-08-04 FR FR7823116A patent/FR2400618A1/en not_active Withdrawn
- 1978-08-11 IT IT26716/78A patent/IT1098088B/en active
- 1978-08-16 DE DE19782835903 patent/DE2835903A1/en not_active Withdrawn
- 1978-08-18 JP JP10009078A patent/JPS5452216A/en active Pending
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3034298A (en) * | 1958-06-12 | 1962-05-15 | Gen Motors Corp | Turbine cooling system |
US3224194A (en) * | 1963-06-26 | 1965-12-21 | Curtiss Wright Corp | Gas turbine engine |
US3528250A (en) * | 1969-04-16 | 1970-09-15 | Gen Motors Corp | Bypass engine with afterburning and compressor bleed air heat exchanger in bypass duct |
US3651645A (en) * | 1969-10-11 | 1972-03-28 | Mtu Muenchen Gmbh | Gas turbine for aircrafts |
US3584458A (en) * | 1969-11-25 | 1971-06-15 | Gen Motors Corp | Turbine cooling |
US3797561A (en) * | 1970-10-02 | 1974-03-19 | Secr Defence | Oil tanks and coolers |
US3842597A (en) * | 1973-03-16 | 1974-10-22 | Gen Electric | Gas turbine engine with means for reducing the formation and emission of nitrogen oxides |
Cited By (273)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4709545A (en) * | 1983-05-31 | 1987-12-01 | United Technologies Corporation | Bearing compartment protection system |
US4546605A (en) * | 1983-12-16 | 1985-10-15 | United Technologies Corporation | Heat exchange system |
US4561246A (en) * | 1983-12-23 | 1985-12-31 | United Technologies Corporation | Bearing compartment for a gas turbine engine |
US4574584A (en) * | 1983-12-23 | 1986-03-11 | United Technologies Corporation | Method of operation for a gas turbine engine |
US4542623A (en) * | 1983-12-23 | 1985-09-24 | United Technologies Corporation | Air cooler for providing buffer air to a bearing compartment |
US4601202A (en) * | 1983-12-27 | 1986-07-22 | General Electric Company | Gas turbine engine component cooling system |
US4608819A (en) * | 1983-12-27 | 1986-09-02 | General Electric Company | Gas turbine engine component cooling system |
US4791782A (en) * | 1986-08-27 | 1988-12-20 | Rolls-Royce Plc | Fluid outlet duct |
WO1990001624A1 (en) * | 1988-08-09 | 1990-02-22 | Sundstrand Corporation | High pressure intercooled turbine engine |
US5038560A (en) * | 1988-10-22 | 1991-08-13 | Rolls-Royce Plc | Fluid outlet duct |
US5012646A (en) * | 1988-11-28 | 1991-05-07 | Machen, Inc. | Turbine engine having combustor air precooler |
US4991394A (en) * | 1989-04-03 | 1991-02-12 | Allied-Signal Inc. | High performance turbine engine |
US5163285A (en) * | 1989-12-28 | 1992-11-17 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." | Cooling system for a gas turbine |
US5203163A (en) * | 1990-08-01 | 1993-04-20 | General Electric Company | Heat exchange arrangement in a gas turbine engine fan duct for cooling hot bleed air |
US5269133A (en) * | 1991-06-18 | 1993-12-14 | General Electric Company | Heat exchanger for cooling a gas turbine |
US5269135A (en) * | 1991-10-28 | 1993-12-14 | General Electric Company | Gas turbine engine fan cooled heat exchanger |
EP0564135A2 (en) * | 1992-03-23 | 1993-10-06 | General Electric Company | Gas turbine engine cooling system |
EP0564135A3 (en) * | 1992-03-23 | 1994-03-23 | Gen Electric | |
US5392614A (en) * | 1992-03-23 | 1995-02-28 | General Electric Company | Gas turbine engine cooling system |
US5319927A (en) * | 1992-07-18 | 1994-06-14 | Rolls-Royce Plc | Ducted fan gas turbine engine |
US5297386A (en) * | 1992-08-26 | 1994-03-29 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) | Cooling system for a gas turbine engine compressor |
US5394687A (en) * | 1993-12-03 | 1995-03-07 | The United States Of America As Represented By The Department Of Energy | Gas turbine vane cooling system |
US5782077A (en) * | 1995-05-15 | 1998-07-21 | Aerospatiale Societe Nationale Industrielle | Device for bleeding off and cooling hot air in an aircraft engine |
US5581996A (en) * | 1995-08-16 | 1996-12-10 | General Electric Company | Method and apparatus for turbine cooling |
US6106229A (en) * | 1997-12-22 | 2000-08-22 | United Technologies Corporation | Heat exchanger system for a gas turbine engine |
US6058696A (en) * | 1997-12-22 | 2000-05-09 | United Technologies Corporation | Inlet and outlet module for a heat exchanger for a flowpath for working medium gases |
US6134880A (en) * | 1997-12-31 | 2000-10-24 | Concepts Eti, Inc. | Turbine engine with intercooler in bypass air passage |
WO2001065095A1 (en) * | 2000-02-29 | 2001-09-07 | Mtu Aero Engines Gmbh | Cooling air system |
US6612114B1 (en) | 2000-02-29 | 2003-09-02 | Daimlerchrysler Ag | Cooling air system for gas turbine |
US20030147741A1 (en) * | 2000-05-15 | 2003-08-07 | Alessandro Coppola | Device for controlling the cooling flows of gas turbines |
US6767182B2 (en) * | 2000-05-15 | 2004-07-27 | Nuovo Pignone Holding S.P.A. | Device for controlling the cooling flows of gas turbines |
US7000404B2 (en) | 2003-07-28 | 2006-02-21 | Snecma Moteurs | Heat exchanger on a turbine cooling circuit |
EP1503061A1 (en) * | 2003-07-28 | 2005-02-02 | Snecma Moteurs | Method for cooling hot turbine parts using air partly cooled by an external heat exchanger and correspondingly cooled gas turbine engine |
US20050022535A1 (en) * | 2003-07-28 | 2005-02-03 | Snecma Moteurs | Heat exchanger on a turbine cooling circuit |
FR2858358A1 (en) * | 2003-07-28 | 2005-02-04 | Snecma Moteurs | METHOD FOR COOLING, BY COOLED AIR IN PART IN AN EXTERNAL EXCHANGER, HOT PARTS OF A TURBOJET ENGINE AND TURBOREACTOR THUS COOLED |
US20050150970A1 (en) * | 2004-01-13 | 2005-07-14 | Snecma Moteurs | Cooling system for hot parts of an aircraft engine, and aircraft engine equipped with such a cooling system |
US20050268612A1 (en) * | 2004-04-24 | 2005-12-08 | Rolls-Royce Plc | Engine |
US7716913B2 (en) * | 2004-04-24 | 2010-05-18 | Rolls-Royce Plc | Engine |
US20060042225A1 (en) * | 2004-08-27 | 2006-03-02 | Pratt & Whitney Canada Corp. | Bypass duct fluid cooler |
US7377100B2 (en) * | 2004-08-27 | 2008-05-27 | Pratt & Whitney Canada Corp. | Bypass duct fluid cooler |
EP1650407A3 (en) * | 2004-10-19 | 2011-07-27 | General Electric Company | Method and apparatus for cooling gas turbine engines |
US8250852B2 (en) | 2006-01-19 | 2012-08-28 | Airbus Operations Sas | Dual flow turbine engine equipped with a precooler |
CN101374722B (en) * | 2006-01-19 | 2011-07-06 | 法国空中巴士公司 | Dual flow turbine engine equipped with a precooler |
CN101374723B (en) * | 2006-01-19 | 2011-06-01 | 法国空中巴士公司 | Dual flow turbine engine equipped with a precooler |
US20090007567A1 (en) * | 2006-01-19 | 2009-01-08 | Airbus France | Dual Flow Turbine Engine Equipped with a Precooler |
US7886520B2 (en) * | 2006-04-20 | 2011-02-15 | Rolls-Royce Plc | Gas turbine engine |
US20070245738A1 (en) * | 2006-04-20 | 2007-10-25 | Stretton Richard G | Heat exchanger arrangement |
US20070245739A1 (en) * | 2006-04-20 | 2007-10-25 | Stretton Richard G | Gas turbine engine |
US7810312B2 (en) * | 2006-04-20 | 2010-10-12 | Rolls-Royce Plc | Heat exchanger arrangement |
US8776952B2 (en) * | 2006-05-11 | 2014-07-15 | United Technologies Corporation | Thermal management system for turbofan engines |
US20070264133A1 (en) * | 2006-05-11 | 2007-11-15 | United Technologies Corporation | Thermal management system for turbofan engines |
US20090301057A1 (en) * | 2006-06-27 | 2009-12-10 | Airbus France | Turboreactor for aircraft |
US8739516B2 (en) * | 2006-06-27 | 2014-06-03 | Airbus Operations Sas | Turboreactor for aircraft |
EP2032822B1 (en) * | 2006-06-27 | 2018-05-30 | Airbus Operations (S.A.S) | Turboreactor for aircraft |
US7658060B2 (en) * | 2006-07-19 | 2010-02-09 | United Technologies Corporation | Lubricant cooling exchanger dual intake duct |
US20080016845A1 (en) * | 2006-07-19 | 2008-01-24 | United Technologies Corporation | Lubricant cooling exchanger dual intake duct |
US20080141656A1 (en) * | 2006-08-01 | 2008-06-19 | Snecma | Bypass turbomachine with artificial variation of its throat section |
US7950218B2 (en) * | 2006-08-01 | 2011-05-31 | Snecma | Bypass turbomachine with artificial variation of its throat section |
US7861512B2 (en) | 2006-08-29 | 2011-01-04 | Pratt & Whitney Canada Corp. | Turbofan bypass duct air cooled fluid cooler installation |
US20100278642A1 (en) * | 2006-08-29 | 2010-11-04 | Pratt & Whitney Canada Corp. | Bypass lip seal |
US20080053059A1 (en) * | 2006-08-29 | 2008-03-06 | Pratt & Whitney Canada Corp. | Turbofan bypass duct air cooled fluid cooler installation |
US20080053060A1 (en) * | 2006-08-29 | 2008-03-06 | Pratt & Whitney Canada Corp. | Bypass lip seal |
EP1898069A3 (en) * | 2006-08-29 | 2011-04-27 | Pratt & Whitney Canada Corp. | Turbofan bypass duct air cooled fluid cooler installation |
WO2008025136A1 (en) * | 2006-08-29 | 2008-03-06 | Pratt & Whitney Canada Corp. | Turbofan bypass duct air cooled fluid cooler installation |
US8387362B2 (en) * | 2006-10-19 | 2013-03-05 | Michael Ralph Storage | Method and apparatus for operating gas turbine engine heat exchangers |
US20080095611A1 (en) * | 2006-10-19 | 2008-04-24 | Michael Ralph Storage | Method and apparatus for operating gas turbine engine heat exchangers |
EP1967716A1 (en) | 2007-02-27 | 2008-09-10 | Snecma | Aircraft engine equipped with heat exchange system |
US20080271433A1 (en) * | 2007-05-03 | 2008-11-06 | Pratt & Whitney Canada Corp. | Low profile bleed air cooler |
US7862293B2 (en) | 2007-05-03 | 2011-01-04 | Pratt & Whitney Canada Corp. | Low profile bleed air cooler |
US8056345B2 (en) | 2007-06-13 | 2011-11-15 | United Technologies Corporation | Hybrid cooling of a gas turbine engine |
US8656722B2 (en) | 2007-06-13 | 2014-02-25 | United Technologies Corporation | Hybrid cooling of a gas turbine engine |
US20080310955A1 (en) * | 2007-06-13 | 2008-12-18 | United Technologies Corporation | Hybrid cooling of a gas turbine engine |
ES2322317A1 (en) * | 2007-06-20 | 2009-06-18 | Futur Investment Partners, S.A. | Turbopropulsor aeronautical (Machine-translation by Google Translate, not legally binding) |
US7856824B2 (en) | 2007-06-25 | 2010-12-28 | Honeywell International Inc. | Cooling systems for use on aircraft |
US20100300066A1 (en) * | 2007-06-25 | 2010-12-02 | Airbus Operations (Societe Par Actions Simplifiee) | Turbojet engine for aircraft |
US20080314047A1 (en) * | 2007-06-25 | 2008-12-25 | Honeywell International, Inc. | Cooling systems for use on aircraft |
EP2011988A2 (en) * | 2007-07-06 | 2009-01-07 | General Electric Company | Heat exchanger for a turbine engine |
EP2011988A3 (en) * | 2007-07-06 | 2014-05-07 | General Electric Company | Heat exchanger for a turbine engine |
US20090007570A1 (en) * | 2007-07-06 | 2009-01-08 | Srikanth Ranganathan | Methods and systems for cooling fluid in a turbine engine |
US8763363B2 (en) | 2007-07-06 | 2014-07-01 | General Electric Company | Method and system for cooling fluid in a turbine engine |
US20090097972A1 (en) * | 2007-10-10 | 2009-04-16 | United Technologies Corp. | Gas Turbine Engine Systems and Related Methods Involving Heat Exchange |
US7946806B2 (en) * | 2007-10-10 | 2011-05-24 | United Technologies Corporation | Gas turbine engine systems and related methods involving heat exchange |
US9212623B2 (en) * | 2007-12-26 | 2015-12-15 | United Technologies Corporation | Heat exchanger arrangement for turbine engine |
US10082081B2 (en) | 2007-12-26 | 2018-09-25 | United Technologies Corporation | Heat exchanger arrangement for turbine engine |
US20090169359A1 (en) * | 2007-12-26 | 2009-07-02 | Michael Joseph Murphy | Heat exchanger arrangement for turbine engine |
US8826641B2 (en) * | 2008-01-28 | 2014-09-09 | United Technologies Corporation | Thermal management system integrated pylon |
US20090188232A1 (en) * | 2008-01-28 | 2009-07-30 | Suciu Gabriel L | Thermal management system integrated pylon |
US20100043396A1 (en) * | 2008-08-25 | 2010-02-25 | General Electric Company | Gas turbine engine fan bleed heat exchanger system |
US8266889B2 (en) | 2008-08-25 | 2012-09-18 | General Electric Company | Gas turbine engine fan bleed heat exchanger system |
US20110162387A1 (en) * | 2008-10-03 | 2011-07-07 | Rolls-Royce Plc | Turbine cooling system |
US9303526B2 (en) * | 2008-10-03 | 2016-04-05 | Rolls-Royce Plc | Turbine cooling system |
WO2010037450A2 (en) | 2008-10-03 | 2010-04-08 | Rolls-Royce Plc | Turbine cooling system |
US20100139288A1 (en) * | 2008-12-10 | 2010-06-10 | Pratt & Whitney Canada Corp. | Heat exchanger to cool turbine air cooling flow |
US8181443B2 (en) | 2008-12-10 | 2012-05-22 | Pratt & Whitney Canada Corp. | Heat exchanger to cool turbine air cooling flow |
US8778063B2 (en) | 2009-02-04 | 2014-07-15 | Purdue Research Foundation | Coiled and microchannel heat exchangers for metal hydride storage systems |
US8636836B2 (en) | 2009-02-04 | 2014-01-28 | Purdue Research Foundation | Finned heat exchangers for metal hydride storage systems |
GB2468346B (en) * | 2009-03-06 | 2011-06-22 | Rolls Royce Plc | Cooling system for an aero gas turbine engine |
US8833053B2 (en) | 2009-03-06 | 2014-09-16 | Rolls-Royce Plc | Cooling system for an aero gas turbine engine |
GB2468346A (en) * | 2009-03-06 | 2010-09-08 | Rolls Royce Plc | Cooling system for gas turbine engine |
US20100303616A1 (en) * | 2009-03-06 | 2010-12-02 | Rolls-Royce Plc | Cooling system for an aero gas turbine engine |
US20110011096A1 (en) * | 2009-07-15 | 2011-01-20 | Rolls-Royce Plc | System for cooling cooling-air in a gas turbine engine |
EP2275656A2 (en) | 2009-07-15 | 2011-01-19 | Rolls-Royce plc | System for cooling cooling-air in a gas turbine engine |
US20110011058A1 (en) * | 2009-07-17 | 2011-01-20 | Rolls-Royce Deutschland Ltd & Co Kg | Turbofan engine |
US8307662B2 (en) | 2009-10-15 | 2012-11-13 | General Electric Company | Gas turbine engine temperature modulated cooling flow |
US20110088405A1 (en) * | 2009-10-15 | 2011-04-21 | John Biagio Turco | Gas turbine engine temperature modulated cooling flow |
US8967958B2 (en) * | 2009-12-21 | 2015-03-03 | Techspace Aero S.A. | Integration of an air-liquid heat exchanger on an engine |
US20110150634A1 (en) * | 2009-12-21 | 2011-06-23 | Denis Bajusz | Integration of an Air-Liquid Heat Exchanger on an Engine |
EP2519723A2 (en) * | 2009-12-31 | 2012-11-07 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and heat exchange system |
US8910465B2 (en) * | 2009-12-31 | 2014-12-16 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and heat exchange system |
EP2519723A4 (en) * | 2009-12-31 | 2014-09-24 | Rolls Royce Nam Tech Inc | Gas turbine engine and heat exchange system |
WO2011139317A2 (en) | 2009-12-31 | 2011-11-10 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and heat exchange system |
US20120144843A1 (en) * | 2009-12-31 | 2012-06-14 | Eric Sean Donovan | Gas turbine engine and cooling system |
US8756910B2 (en) * | 2009-12-31 | 2014-06-24 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and cooling system |
US20120144842A1 (en) * | 2009-12-31 | 2012-06-14 | Snyder Douglas J | Gas turbine engine and heat exchange system |
US20120282079A1 (en) * | 2010-04-09 | 2012-11-08 | Glahn Jorn A | Rear hub cooling for high pressure compressor |
US8459040B2 (en) * | 2010-04-09 | 2013-06-11 | United Technologies Corporation | Rear hub cooling for high pressure compressor |
US8266888B2 (en) * | 2010-06-24 | 2012-09-18 | Pratt & Whitney Canada Corp. | Cooler in nacelle with radial coolant |
EP2447543A1 (en) * | 2010-10-27 | 2012-05-02 | Siemens Aktiengesellschaft | Axial compressor and method for operating same |
US8602717B2 (en) | 2010-10-28 | 2013-12-10 | United Technologies Corporation | Compression system for turbomachine heat exchanger |
US8784047B2 (en) * | 2010-11-04 | 2014-07-22 | Hamilton Sundstrand Corporation | Gas turbine engine heat exchanger with tapered fins |
US20120114467A1 (en) * | 2010-11-04 | 2012-05-10 | Elder James S | Gas turbine engine heat exchanger with tapered fins |
US9051943B2 (en) | 2010-11-04 | 2015-06-09 | Hamilton Sundstrand Corporation | Gas turbine engine heat exchanger fins with periodic gaps |
EP2655842A4 (en) * | 2010-12-24 | 2015-07-29 | Rolls Royce Nam Tech Inc | Gas turbine engine heat exchanger |
US9410482B2 (en) * | 2010-12-24 | 2016-08-09 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine heat exchanger |
US20120159961A1 (en) * | 2010-12-24 | 2012-06-28 | Michael Stephen Krautheim | Gas turbine engine heat exchanger |
WO2012088543A1 (en) | 2010-12-24 | 2012-06-28 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine heat exchanger |
US8397487B2 (en) | 2011-02-28 | 2013-03-19 | General Electric Company | Environmental control system supply precooler bypass |
EP2527603A2 (en) | 2011-05-27 | 2012-11-28 | General Electric Company | Flade Duct Turbine Cooling And Power And Thermal Management |
US9045998B2 (en) | 2011-12-12 | 2015-06-02 | Honeywell International Inc. | System for directing air flow to a plurality of plena |
US20130152602A1 (en) * | 2011-12-14 | 2013-06-20 | Rolls-Royce Plc | Controller |
US9255492B2 (en) | 2011-12-14 | 2016-02-09 | Rolls-Royce Plc | Gas turbine engine having a multi-variable closed loop controller for regulating tip clearance |
US9249729B2 (en) * | 2011-12-14 | 2016-02-02 | Rolls-Royce Plc | Turbine component cooling with closed looped control of coolant flow |
US9222411B2 (en) | 2011-12-21 | 2015-12-29 | General Electric Company | Bleed air and hot section component cooling air system and method |
US20140338334A1 (en) * | 2011-12-30 | 2014-11-20 | Rolls-Royce North American Technologies, Inc. | Aircraft propulsion gas turbine engine with heat exchange |
WO2013147953A1 (en) * | 2011-12-30 | 2013-10-03 | Rolls-Royce North American Technologies Inc. | Aircraft propulsion gas turbine engine with heat exchange |
US9771867B2 (en) * | 2011-12-30 | 2017-09-26 | Rolls-Royce Corporation | Gas turbine engine with air/fuel heat exchanger |
US9109514B2 (en) | 2012-01-10 | 2015-08-18 | Hamilton Sundstrand Corporation | Air recovery system for precooler heat-exchanger |
US20130186102A1 (en) * | 2012-01-25 | 2013-07-25 | Honeywell International Inc. | Gas turbine engine in-board cooled cooling air system |
US9243563B2 (en) * | 2012-01-25 | 2016-01-26 | Honeywell International Inc. | Gas turbine engine in-board cooled cooling air system |
US9267434B2 (en) | 2012-01-29 | 2016-02-23 | United Technologies Corporation | Heat exchanger |
US20130239583A1 (en) * | 2012-03-14 | 2013-09-19 | United Technologies Corporation | Pump system for hpc eps parasitic loss elimination |
US9394803B2 (en) * | 2012-03-14 | 2016-07-19 | United Technologies Corporation | Bypass air-pump system within the core engine to provide air for an environmental control system in a gas turbine engine |
US9267390B2 (en) | 2012-03-22 | 2016-02-23 | Honeywell International Inc. | Bi-metallic actuator for selectively controlling air flow between plena in a gas turbine engine |
US10072577B2 (en) | 2012-05-02 | 2018-09-11 | Pratt & Whitney Canada Corp. | Air cooler system for gas turbine engines |
US9388739B2 (en) * | 2012-05-02 | 2016-07-12 | Pratt & Whitney Canada Corp. | Air cooler system for gas turbine engines |
US20130291554A1 (en) * | 2012-05-02 | 2013-11-07 | Pratt & Whitney Canada Corp. | Air cooler system for gas turbine engines |
US9945252B2 (en) * | 2012-07-05 | 2018-04-17 | United Technologies Corporation | Gas turbine engine oil tank with integrated packaging configuration |
US20140010639A1 (en) * | 2012-07-05 | 2014-01-09 | Nathan Snape | Gas turbine engine oil tank with integrated packaging configuration |
US10641128B2 (en) * | 2012-07-05 | 2020-05-05 | United Technologies Corporation | Gas turbine engine oil tank with integrated packaging configuration |
US20150192033A1 (en) * | 2012-07-19 | 2015-07-09 | Snecma | Cooling of an oil circuit of a turbomachine |
US10352190B2 (en) * | 2012-07-19 | 2019-07-16 | Safran Aircraft Engines | Cooling of an oil circuit of a turbomachine |
US20150247462A1 (en) * | 2012-09-28 | 2015-09-03 | United Technologies Corporation | Gas turbine engine thermal management system for heat exchanger using bypass flow |
US10036329B2 (en) * | 2012-09-28 | 2018-07-31 | United Technologies Corporation | Gas turbine engine thermal management system for heat exchanger using bypass flow |
US10385777B2 (en) * | 2012-10-01 | 2019-08-20 | United Technologies Corporation | Bifurcated inlet scoop for gas turbine engine |
US20140096534A1 (en) * | 2012-10-04 | 2014-04-10 | United Technologies Corporation | Low Profile Compressor Bleed Air-Oil Coolers |
US9714610B2 (en) * | 2012-10-04 | 2017-07-25 | United Technologies Corporation | Low profile compressor bleed air-oil coolers |
US20140131027A1 (en) * | 2012-11-09 | 2014-05-15 | Rolls-Royce Plc | Heat exchange arrangement |
US9562475B2 (en) | 2012-12-19 | 2017-02-07 | Siemens Aktiengesellschaft | Vane carrier temperature control system in a gas turbine engine |
US20150128561A1 (en) * | 2013-03-13 | 2015-05-14 | Rolls-Royce North American Technologies, Inc. | Three stream, variable area, vectorable nozzle |
US9845768B2 (en) * | 2013-03-13 | 2017-12-19 | Rolls-Royce North American Technologies, Inc. | Three stream, variable area, vectorable nozzle |
US9856746B2 (en) * | 2013-03-14 | 2018-01-02 | United Technologies Corporation | Heatshield discourager seal for a gas turbine engine |
US20160032763A1 (en) * | 2013-03-14 | 2016-02-04 | United Technologies Corporation | Heatshield discourager seal for a gas turbine engine |
US10352191B2 (en) * | 2013-03-15 | 2019-07-16 | United Technologies Corporation | Gas turbine engine with air-oil cooler oil tank |
US20160024964A1 (en) * | 2013-03-15 | 2016-01-28 | United Technologies Corporation | Gas turbine engine with air-oil cooler oil tank |
US9429072B2 (en) | 2013-05-22 | 2016-08-30 | General Electric Company | Return fluid air cooler system for turbine cooling with optional power extraction |
US9422063B2 (en) | 2013-05-31 | 2016-08-23 | General Electric Company | Cooled cooling air system for a gas turbine |
EP3049641A4 (en) * | 2013-09-24 | 2017-06-28 | United Technologies Corporation | Bypass duct heat exchanger placement |
WO2015047533A1 (en) | 2013-09-24 | 2015-04-02 | United Technologies Corporation | Bypass duct heat exchanger placement |
RU2550224C1 (en) * | 2013-11-25 | 2015-05-10 | Российская Федерация, от имени которой выступает Министерство промышленности и торговли Российской Федерации (Минпромторг России) | Gas turbine engine |
WO2015105552A1 (en) | 2014-01-07 | 2015-07-16 | United Technologies Corporation | Cross-stream heat exchanger |
US10287983B2 (en) * | 2014-01-07 | 2019-05-14 | United Technologies Corporation | Cross-stream heat exchanger |
US20170260905A1 (en) * | 2014-01-07 | 2017-09-14 | United Technologies Corporation | Cross-stream heat exchanger |
US10907546B2 (en) | 2014-01-07 | 2021-02-02 | Raytheon Technologies Corporation | Cross-stream heat exchanger |
US20150300254A1 (en) * | 2014-04-17 | 2015-10-22 | Rolls-Royce Plc | Propulsion engine |
US10066550B2 (en) | 2014-05-15 | 2018-09-04 | Rolls-Royce North American Technologies, Inc. | Fan by-pass duct for intercooled turbo fan engines |
US20150354465A1 (en) * | 2014-06-06 | 2015-12-10 | United Technologies Corporation | Turbine stage cooling |
EP2952698A1 (en) * | 2014-06-06 | 2015-12-09 | United Technologies Corporation | Turbine stage cooling |
US11684974B2 (en) | 2014-10-21 | 2023-06-27 | Raytheon Technologies Corporation | Additive manufactured ducted heat exchanger system |
US20160230658A1 (en) * | 2015-02-10 | 2016-08-11 | United Technologies Corporation | Gas turbine engine and an airflow control system |
US11236639B2 (en) * | 2015-02-10 | 2022-02-01 | Raytheon Technologies Corporation | Gas turbine engine and an airflow control system |
US20170082028A1 (en) * | 2015-02-12 | 2017-03-23 | United Technologies Corporation | Intercooled cooling air using existing heat exchanger |
US10830149B2 (en) | 2015-02-12 | 2020-11-10 | Raytheon Technologies Corporation | Intercooled cooling air using cooling compressor as starter |
US10731560B2 (en) * | 2015-02-12 | 2020-08-04 | Raytheon Technologies Corporation | Intercooled cooling air |
US20160237905A1 (en) * | 2015-02-12 | 2016-08-18 | United Technologies Corporation | Intercooled cooling air |
US20160237908A1 (en) * | 2015-02-12 | 2016-08-18 | United Technologies Corporation | Intercooled cooling air using existing heat exchanger |
US20160237907A1 (en) * | 2015-02-12 | 2016-08-18 | United Technologies Corporation | Intercooled cooling air with auxiliary compressor control |
US11808210B2 (en) | 2015-02-12 | 2023-11-07 | Rtx Corporation | Intercooled cooling air with heat exchanger packaging |
US10371055B2 (en) | 2015-02-12 | 2019-08-06 | United Technologies Corporation | Intercooled cooling air using cooling compressor as starter |
US10480419B2 (en) | 2015-04-24 | 2019-11-19 | United Technologies Corporation | Intercooled cooling air with plural heat exchangers |
US10221862B2 (en) | 2015-04-24 | 2019-03-05 | United Technologies Corporation | Intercooled cooling air tapped from plural locations |
US10830148B2 (en) | 2015-04-24 | 2020-11-10 | Raytheon Technologies Corporation | Intercooled cooling air with dual pass heat exchanger |
US11215197B2 (en) | 2015-04-24 | 2022-01-04 | Raytheon Technologies Corporation | Intercooled cooling air tapped from plural locations |
US10100739B2 (en) | 2015-05-18 | 2018-10-16 | United Technologies Corporation | Cooled cooling air system for a gas turbine engine |
US10914235B2 (en) | 2015-05-18 | 2021-02-09 | Raytheon Technologies Corporation | Cooled cooling air system for a gas turbine engine |
US10794288B2 (en) | 2015-07-07 | 2020-10-06 | Raytheon Technologies Corporation | Cooled cooling air system for a turbofan engine |
US10677166B2 (en) | 2015-08-12 | 2020-06-09 | Rolls-Royce North American Technologies Inc. | Heat exchanger for a gas turbine engine propulsion system |
US10443508B2 (en) | 2015-12-14 | 2019-10-15 | United Technologies Corporation | Intercooled cooling air with auxiliary compressor control |
US11512651B2 (en) | 2015-12-14 | 2022-11-29 | Raytheon Technologies Corporation | Intercooled cooling air with auxiliary compressor control |
US11002195B2 (en) | 2015-12-14 | 2021-05-11 | Raytheon Technologies Corporation | Intercooled cooling air with auxiliary compressor control |
US20170167377A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam Via Turbine Extraction |
US20170167379A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam via Turbine Extraction and Compressor Extraction |
US10072573B2 (en) * | 2015-12-15 | 2018-09-11 | General Electric Company | Power plant including an ejector and steam generating system via turbine extraction |
US9976479B2 (en) * | 2015-12-15 | 2018-05-22 | General Electric Company | Power plant including a static mixer and steam generating system via turbine extraction and compressor extraction |
US9970354B2 (en) * | 2015-12-15 | 2018-05-15 | General Electric Company | Power plant including an ejector and steam generating system via turbine extraction and compressor extraction |
US20170167378A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam Via Turbine Extraction and Compressor Extraction |
US9964035B2 (en) * | 2015-12-15 | 2018-05-08 | General Electric Company | Power plant including exhaust gas coolant injection system and steam generating system via turbine extraction |
US20170167376A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | System for Generating Steam Via Turbine Extraction |
US9890710B2 (en) * | 2015-12-15 | 2018-02-13 | General Electric Company | Power plant with steam generation via combustor gas extraction |
US20170167375A1 (en) * | 2015-12-15 | 2017-06-15 | General Electric Company | Power Plant With Steam Generation Via Combustor Gas Extraction |
US10125684B2 (en) | 2015-12-29 | 2018-11-13 | Pratt & Whitney Canada Corp. | Surface cooler for aero engine |
US20170184027A1 (en) * | 2015-12-29 | 2017-06-29 | General Electric Company | Method and system for compressor and turbine cooling |
US20170307311A1 (en) * | 2016-04-26 | 2017-10-26 | United Technologies Corporation | Simple Heat Exchanger Using Super Alloy Materials for Challenging Applications |
RU2623854C1 (en) * | 2016-07-06 | 2017-06-29 | Публичное акционерное общество "Научно-производственное объединение "Сатурн" | Method of greasing and cooling front support of the rotor of the gas turbine engine |
US20180051630A1 (en) * | 2016-08-22 | 2018-02-22 | United Technologies Corporation | Heat Exchanger for Gas Turbine Engine with Support Damper Mounting |
US10436115B2 (en) * | 2016-08-22 | 2019-10-08 | United Technologies Corporation | Heat exchanger for gas turbine engine with support damper mounting |
US10669940B2 (en) | 2016-09-19 | 2020-06-02 | Raytheon Technologies Corporation | Gas turbine engine with intercooled cooling air and turbine drive |
US11236675B2 (en) | 2016-09-19 | 2022-02-01 | Raytheon Technologies Corporation | Gas turbine engine with intercooled cooling air and turbine drive |
US10794290B2 (en) | 2016-11-08 | 2020-10-06 | Raytheon Technologies Corporation | Intercooled cooled cooling integrated air cycle machine |
US10550768B2 (en) | 2016-11-08 | 2020-02-04 | United Technologies Corporation | Intercooled cooled cooling integrated air cycle machine |
US11066997B2 (en) * | 2016-12-14 | 2021-07-20 | Safran Aircraft Engines | Fluid circuit in a turbine engine |
US10961911B2 (en) | 2017-01-17 | 2021-03-30 | Raytheon Technologies Corporation | Injection cooled cooling air system for a gas turbine engine |
US11846237B2 (en) | 2017-01-19 | 2023-12-19 | Rtx Corporation | Gas turbine engine with intercooled cooling air and dual towershaft accessory gearbox |
US10995673B2 (en) | 2017-01-19 | 2021-05-04 | Raytheon Technologies Corporation | Gas turbine engine with intercooled cooling air and dual towershaft accessory gearbox |
US10837364B2 (en) * | 2017-01-27 | 2020-11-17 | Raytheon Technologies Corporation | Thermal shield for gas turbine engine diffuser case |
US20190003392A1 (en) * | 2017-01-27 | 2019-01-03 | United Technologies Corporation | Thermal shield for gas turbine engine diffuser case |
US10577964B2 (en) | 2017-03-31 | 2020-03-03 | United Technologies Corporation | Cooled cooling air for blade air seal through outer chamber |
US11773742B2 (en) | 2017-03-31 | 2023-10-03 | Rtx Corporation | Cooled cooling air for blade air seal through outer chamber |
US10711640B2 (en) | 2017-04-11 | 2020-07-14 | Raytheon Technologies Corporation | Cooled cooling air to blade outer air seal passing through a static vane |
US11268444B2 (en) * | 2017-05-18 | 2022-03-08 | Raytheon Technologies Corporation | Turbine cooling arrangement |
CN109723558A (en) * | 2017-10-30 | 2019-05-07 | 通用电气公司 | Gas-turbine unit and its operating method including heat management system |
CN109723558B (en) * | 2017-10-30 | 2022-07-29 | 通用电气公司 | Gas turbine engine including thermal management system and method of operating the same |
US11174790B2 (en) * | 2017-11-14 | 2021-11-16 | Rolls-Royce Plc | Gas turbine engine having an air-oil heat exchanger |
US20190145317A1 (en) * | 2017-11-14 | 2019-05-16 | Rolls-Royce Plc | Gas turbine engine having an air-oil heat exchanger |
US11542026B2 (en) * | 2017-12-21 | 2023-01-03 | Safran Nacelles | Aircraft propulsion unit and method for ventilating an engine enclosure |
US11725584B2 (en) | 2018-01-17 | 2023-08-15 | General Electric Company | Heat engine with heat exchanger |
RU2679573C1 (en) * | 2018-02-16 | 2019-02-11 | Валерий Николаевич Сиротин | Cooling system of bearings of gas turbine engine turbines |
US10738703B2 (en) | 2018-03-22 | 2020-08-11 | Raytheon Technologies Corporation | Intercooled cooling air with combined features |
US20190301301A1 (en) * | 2018-04-02 | 2019-10-03 | General Electric Company | Cooling structure for a turbomachinery component |
US10808572B2 (en) * | 2018-04-02 | 2020-10-20 | General Electric Company | Cooling structure for a turbomachinery component |
US10808619B2 (en) * | 2018-04-19 | 2020-10-20 | Raytheon Technologies Corporation | Intercooled cooling air with advanced cooling system |
US10830145B2 (en) | 2018-04-19 | 2020-11-10 | Raytheon Technologies Corporation | Intercooled cooling air fleet management system |
US20190323431A1 (en) * | 2018-04-19 | 2019-10-24 | United Technologies Corporation | Intercooled cooling air with advanced cooling system |
US11162417B2 (en) * | 2018-05-22 | 2021-11-02 | General Electric Company | Scoop inlet |
RU187493U1 (en) * | 2018-05-28 | 2019-03-11 | Публичное Акционерное Общество "Одк-Сатурн" | HEAT EXCHANGER COOLING DEVICE |
US10718233B2 (en) | 2018-06-19 | 2020-07-21 | Raytheon Technologies Corporation | Intercooled cooling air with low temperature bearing compartment air |
US11060484B2 (en) * | 2018-06-29 | 2021-07-13 | The Boeing Company | Nozzle wall for an air-breathing engine of a vehicle and method therefor |
US11255268B2 (en) | 2018-07-31 | 2022-02-22 | Raytheon Technologies Corporation | Intercooled cooling air with selective pressure dump |
US11773780B2 (en) | 2018-07-31 | 2023-10-03 | Rtx Corporation | Intercooled cooling air with selective pressure dump |
US11300002B2 (en) | 2018-12-07 | 2022-04-12 | Pratt & Whitney Canada Corp. | Static take-off port |
US11549434B2 (en) | 2019-02-06 | 2023-01-10 | Raytheon Technologies Corporation | Engine bleed air ducting into heat exchanger |
US11078837B2 (en) * | 2019-02-06 | 2021-08-03 | Raytheon Technologies Corporation | Engine bleed air ducting into heat exchanger |
US11549393B2 (en) * | 2019-02-18 | 2023-01-10 | Safran Aero Boosters Sa | Air-oil heat exchanger |
US11174816B2 (en) | 2019-02-25 | 2021-11-16 | Rolls-Royce Corporation | Bypass duct conformal heat exchanger array |
WO2020234525A3 (en) * | 2019-05-20 | 2021-01-14 | Safran | Optimised heat exchange system of a turbomachine |
US11655761B2 (en) * | 2019-05-20 | 2023-05-23 | Safran | Optimized heat exchange system for a turbomachine |
WO2020234524A1 (en) | 2019-05-20 | 2020-11-26 | Safran | Optimized heat exchange system for a turbomachine |
CN113966433A (en) * | 2019-05-20 | 2022-01-21 | 赛峰集团 | Optimized heat exchange system for a turbomachine |
WO2020234525A2 (en) | 2019-05-20 | 2020-11-26 | Safran | Optimised heat exchange system |
FR3096444A1 (en) * | 2019-05-20 | 2020-11-27 | Safran | OPTIMIZED HEAT EXCHANGE SYSTEM |
FR3096409A1 (en) * | 2019-05-20 | 2020-11-27 | Safran | OPTIMIZED HEAT EXCHANGE SYSTEM |
US11891955B2 (en) | 2019-05-20 | 2024-02-06 | Safran | Optimised heat exchange system of a turbomachine |
US11492971B2 (en) * | 2019-09-06 | 2022-11-08 | Raytheon Technologies Corporation | Turbine engine system with heat exchanger in bypassable secondary duct |
CN111577466A (en) * | 2020-06-22 | 2020-08-25 | 中国航空发动机研究院 | Ice-proof bleed air preheating and turbine cooling bleed air precooling system for aircraft engine |
CN112228226A (en) * | 2020-10-16 | 2021-01-15 | 中国航发四川燃气涡轮研究院 | Aircraft engine turbine rotor cooling thermal management system |
US11788469B2 (en) | 2020-11-16 | 2023-10-17 | General Electric Company | Thermal management system for a gas turbine engine |
US11512639B2 (en) * | 2021-01-26 | 2022-11-29 | General Electric Company | Heat transfer system |
US20220235705A1 (en) * | 2021-01-26 | 2022-07-28 | General Electric Company | Heat transfer system |
US11788470B2 (en) | 2021-03-01 | 2023-10-17 | General Electric Company | Gas turbine engine thermal management |
US11591965B2 (en) | 2021-03-29 | 2023-02-28 | General Electric Company | Thermal management system for transferring heat between fluids |
US11674396B2 (en) | 2021-07-30 | 2023-06-13 | General Electric Company | Cooling air delivery assembly |
US11702958B2 (en) | 2021-09-23 | 2023-07-18 | General Electric Company | System and method of regulating thermal transport bus pressure |
CN114876644A (en) * | 2022-05-09 | 2022-08-09 | 北京航空航天大学 | Periodic porous bearing support plate |
CN114876644B (en) * | 2022-05-09 | 2024-01-19 | 北京航空航天大学 | Periodic porous bearing support plate |
Also Published As
Publication number | Publication date |
---|---|
FR2400618A1 (en) | 1979-03-16 |
IT1098088B (en) | 1985-08-31 |
IT7826716A0 (en) | 1978-08-11 |
JPS5452216A (en) | 1979-04-24 |
DE2835903A1 (en) | 1979-03-01 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US4254618A (en) | Cooling air cooler for a gas turbofan engine | |
CA1095271A (en) | Cooling air cooler for a gas turbine engine | |
US5317877A (en) | Intercooled turbine blade cooling air feed system | |
CA2299148C (en) | Compressor system and methods for reducing cooling airflow | |
US5724806A (en) | Extracted, cooled, compressed/intercooled, cooling/combustion air for a gas turbine engine | |
CN109723558B (en) | Gas turbine engine including thermal management system and method of operating the same | |
CN110529256B (en) | Air cycle assembly for a gas turbine engine assembly | |
US6430931B1 (en) | Gas turbine in-line intercooler | |
US4466239A (en) | Gas turbine engine with improved air cooling circuit | |
EP0564135B1 (en) | Gas turbine engine cooling system | |
US2588532A (en) | Jet propulsion unit | |
US3651645A (en) | Gas turbine for aircrafts | |
CA2042220A1 (en) | Aft entry cooling system and method for an aircraft engine | |
JP2013543556A (en) | Aircraft engine system and method for operating the same | |
WO2002038938A1 (en) | Bypass gas turbine engine and cooling method for working fluid | |
US10989411B2 (en) | Heat exchanger for turbo machine | |
CA2991449A1 (en) | Heat exchanger assembly for engine bleed air | |
US4302148A (en) | Gas turbine engine having a cooled turbine | |
US2721445A (en) | Aircraft propulsion plant of the propeller-jet turbine type | |
US9341119B2 (en) | Cooling air system for aircraft turbine engine | |
CN111255568A (en) | Gas turbine engine | |
CA1194803A (en) | Cool tip combustor | |
US5224819A (en) | Cooling air pick up | |
US6397577B1 (en) | Shaftless gas turbine engine spool | |
CA1105276A (en) | Cooling air cooler for a gas tubofan engine |